Mounting device for an aircraft

ABSTRACT

The invention relates to a mounting device in an aircraft. In one aspect of the invention, a method is disclosed for assembling a metal securement member of the mounting device. In another aspect of the invention, a mounting device is disclosed. In another aspect of the invention, a method is disclosed for installing and using a mounting device in an aircraft.

BACKGROUND OF THE INVENTION

There are a variety of mounting devices and methods of using mountingdevices in the art of manufacturing aircraft. Some of these mountingdevices when installed in aircraft wings cause imperfections on thesurface of the aircraft wing which can increase undesirable turbulentflow and drag over the wing. In addition, some of the mounting devicesused in the art do not contribute enough stiffness and strength toaircraft wings and fuselage. Further, some of the mounting devices donot adequately secure structures within aircraft. A mounting device, andmethod for its use, is needed which may solve one or more problems inone or more of the existing mounting devices used in aircraft.

SUMMARY OF THE INVENTION

In one aspect of the invention, a method is disclosed for assembling ametal securement member. The method comprises holding at lest two metalsheets in a predetermined position, welding the sheets together with thesheets held in the predetermined position, and annealing the sheets withmaintaining the sheets held in the predetermined position. A clampingassembly which can be used with the method for assembling the metalsecurement member is also disclosed.

In another aspect of the invention, a mounting device is disclosed. Themounting device comprises a metal securement member and a compositemember attached to the metal securement member and positioned within achannel formed in a structural member of an aircraft. The metalsecurement member can be configured in the form of at least one of a pichord, T chord, J chord, I beam, sinewave beam, and F beam. The metalsecurement member can be constructed of titanium. The composite membercan be constructed of a graphite fiber epoxy laminate.

One example of the mounting device comprises a titanium alloy securementmember comprising a cap member and two spaced apart arms connected toand extending from a first surface of the cap member and a graphitefiber epoxy laminate member attached to a second surface of the capmember, opposite the first, and positioned within a channel formed in awing of an aircraft.

In another aspect of the invention, a method is disclosed for installinga mounting device in a structural member of an aircraft. The methodcomprising the steps of positioning a mounting device within a channelformed in the structural member, wherein the mounting device comprises ametal securement member and a composite member attached to the metalsecurement member, and attaching the composite member to the structuralmember within the channel.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a top perspective view of an embodiment of the metalsecurement member in the configuration of a pi chord;

FIG. 2 is an end elevational view of an embodiment of the mountingdevice of the present invention;

FIG. 3 is a partially cut away top perspective view of a wing of anaircraft with an embodiment of the mounting device of the presentinvention mounted thereto;

FIG. 4 is an end elevational view of FIG. 3 with a spar engaged to themounting device of the present invention.

FIG. 5 is an enlarged view of a portion of FIG. 4;

FIG. 6 is a top plan view of an aircraft wing being installed on anaircraft with an embodiment of the mounting device of the presentinvention installed along the length of the aircraft wing.

FIG. 7 is an end elevational view of an embodiment of the mountingdevice of the present invention positioned within a channel formed in astructural member of an aircraft;

FIG. 8 is an end elevational view of an embodiment of the mountingdevice of the present invention positioned within a channel formed in astructural member of an aircraft.

FIG. 9 is an end elevational view of an embodiment of the constraint andfixture tool used to construct an embodiment of the metal securementmember in the configuration of a pi chord;

FIG. 10 is a top perspective view of FIG. 9.

DETAILED DESCRIPTION OF THE INVENTION

The following detailed description is of the best currently contemplatedmodes of carrying out the invention. The description is not to be takenin a limiting sense, but is made merely for the purpose of illustratingthe general principles of the invention, since the scope of theinvention is best defined by the appended claims.

As shown in FIG. 1, in one example of the invention, a metal securementmember can be configured in the form of a pi chord (1). Pi chord (1) isconstructed of a cap (2) and two spaced apart arms (3) secured to andextending from the same side of the cap (2). Pi chord (1) generallyresembles the shape of the Greek letter π (pi). Pi chord (1) in thisexample provides a pinch angle (4) where spaced apart arms (3) pinch orconverge toward one another as they extend from cap (2) at an anglerelative to cap (2). The space between spaced apart arms (3) form a slot(5). Cap (2) and spaced apart arms (3) in this example are constructedof three sheets of metal. The metal can be a titanium alloy.

In one example the titanium alloy can be a alpha-beta titanium alloy,such as, for example, Ti-6 Al-4V or Ti 3Al-2.5 V, or a precipitationhardened metastable beta titanium alloy, such as, for example, Ti15-3-3-3, Beta 21S, or VT-22. A particularly useful example ofmetastable beta titanium alloy for use in this invention posses theproperties of high yield strength of approximately 180 to 200 Ksi, amodulus of approximately 14.5 to 15 Msi, and a coefficient of thermalexpansion of approximately 5.1×10−6/degree Fahrenheit. The titaniumalloy sheets can be between 0.020″ to 0.20″ thick.

In one example of the invention, the metal securement member can be madeby using a novel method of assembly disclosed herein. For example, pichord (1) can be manufactured using this method by holding spaced apartarms (3) in contact with cap (2) throughout the welding and annealingsteps. Holding spaced apart arms (3) to cap (2) throughout both thewelding and annealing steps prevents pi chord (1) from distorting inconfiguration. Distortion can occur if the components were otherwisereleased between the welding and annealing steps. The holding of spacedapart arms (3) to cap (2) can be performed by a constraint and fixturetool. An example a constraint and fixture tool (21) used to make a metalsecurement member in the configuration of a pi chord is shown in FIGS. 9and 10. The welding can be performed by utilizing laser, plasma arc, orgas tungsten arc stake welding. The annealing can take place in a vacuumfurnace. This method can be used to make a welded joint between anynumber of metal sheets, including but not limited to metal sheets madeof titanium alloy, aluminum, steel, stainless steel, and nickel basedalloys. This method can be used to fabricate metal securement members ofdifferent shapes, including but not limited to T and J chords; I andsinewave beams; and F beams in either straight or curved sections.

FIGS. 9 and 10 show an example of a constraint and fixture tool (21)used to construct a metal securement member in the configuration of a pichord (1). Three sheets of metal are held in the configuration of a pichord by using a block (22), wedges (24 a and 24 b), an insert (26), andcap clamps (28 a and 28 b), as shown in FIGS. 9 and 10. Clamps (32 a and32 b) and spacers (34 a and 34 b) are used to constrain and fix the toolso that the spaced apart arms (3) and the cap (2) are configured andheld in a desired position. The cap clamps (28 a and 28 b) areconfigured so that a weld trough (30) is defined where the spaced apartarms (3) contact the cap (2) to allow access for welding. The constraintfixture tool (21) of FIGS. 9 and 10 can be used to construct pi chords(1) with slots (5) of various width by changing the width of the insert(26). As the width of the insert (26) increases, the wedges (24 a and 24b) will set at a higher position relative to the bottom of the block(22). This design allows the constraint and fixture tool (21) to be usedto construct pi chords with different slot widths without having tochange all the components of the fixture tool (21).

In one example, this method is employed in the construction of a metalsecurement members made of titanium alloy. When using this method in theconstruction of a metal securement member made of titanium alloy, theannealing step can be performed in a vacuum furnace at about 900 to 1000degrees Fahrenheit for one-half to four hours for metastable betatitanium alloys or a about 1400 to 1450 degrees Fahrenheit for 15 to 30minutes for alpha-beta titanium alloys. Benefits of the annealing stepinclude, but are not limited to stress relief and precipitationstrengthening for the metastable beta titanium alloys and residualstress relief for the alpha-beta titanium alloys.

In another example, holes (8) are defined in the spaced apart arms (3).Holes (8) can be positioned closer to the cap than the distal end of anarm (3 a) and can be positioned along the length of pi chord (1). Aswill be described below, holes (8) can be utilized to permit adhesive(18) to pass from slot (5) or inside arms (3) to outside arms (3).

As shown in FIG. 2, in one example of mounting device (6) a compositemember (7) is bonded to cap (2) of pi chord (1). The composite membercan be constructed of a graphite composite material using thermosetresin or a thermoplastic resin matrix. Suitable thermoset resinsinclude, but are not limited to, epoxy, bismaleimide (“BMI”), andpolyimide (“PI”). Suitable thermoplastic resins include, but are notlimited to, polyetheretherketone (“PEEK”), polyetherketoneketone(“PEKK”), polyetherimide (“PEI”), and polyphenylene sulfide (“PPS”). Thegraphite composite material can have a lay-up of approximately 70 to 80%zero degree oriented fibers and approximately 20 to 30% forty fivedegree oriented fibers. The graphite fiber used to construct thegraphite composite material is preferably a polyacrylonitrile (“PAN”)derived fiber with an elastic modulus of 44 Msi (IM-7 fiber) to 65 Msi(M46J fiber) with a 1.2 to 1.5% strain to failure.

As mentioned above, pi chord (1) can be made of a titanium alloy. Onebenefit of a mounting device (6) with a pi chord (1) constructed oftitanium alloy includes but is not limited to a high shear strength atthe joint where spaced apart arms (3) contact cap (2). The greater shearstrength at this joint provides many benefits, including but not limitedto allowing pi chord (1) to be stiffer and thinner than pi chordsconstructed of other metals.

In another example of mounting device (6), pi chord (1) is constructedof a titanium alloy and composite member (7) is constructed of agraphite fiber epoxy laminate composite material that utilizes PANderived graphite fibers. One benefit of this example is that thetitanium alloy and the graphite fiber epoxy laminate composite havesimilar coefficients of thermal expansion, thereby reducing the amountof internal thermal stresses that can result in either the warping ofmounting device (6) or separation of composite member (7) from pi chord(1). Another benefit of this example is that the high strain to yield oftitanium alloys, particularly metastable beta titanium alloys,correspond to the strain to failure of PAN derived graphite fibers,thereby creating efficient structures.

In the example shown in FIG. 2, a composite member (7) that is precuredcan be bonded to cap (2) using a bonding agent, other adhesive, or both(18). The bonding agent can be a SOL-GEL as disclosed in U.S. Pat. Nos.6,797,376; 6,770,371; 6,037,060; 5,869140; 5,814,137 and relatedpatents. The other adhesive can be, for example, a polymeric filmadhesive (such as FM-300), a paste adhesive, a thermoplastic resin film,or other suitable adhesives used by those skilled in the art. Compositemember (7) can be coextensive with the surface of cap (2) or extendbeyond the perimeter of cap (2). Composite member (7) can also extendbeyond the perimeter of cap (2) and wrap (17) around it to come incontact with the under side of cap (2). Should wrapping of compositemember (7) around cap (2) be employed, the wrap (17) structure preventscomposite member (7) from separating from cap (2) if the bond betweenthe two were to fail. All or a portion of mounting device (6) withcomposite member (7) bonded to cap (2) can be vacuum bagged and cured inan oven at about atmospheric pressure or in an autoclave at greater thanatmospheric pressure, preferably at about 80 to 200 psia.

In another example, composite member (7) can be co-cured directly ontocap (2) of pi chord (1). A bonding agent, other adhesive, or both can beapplied to the surface of cap (2) to facilitate bonding of compositemember to cap (2) during curing. In one example, composite member (7)can be constructed of composite plies that can be lay-ed up on the cap,thereby using it as the lay-up mandrel tool. Composite member (7) can becoextensive with the surface of cap (2) or extend beyond the perimeterof cap (2). Composite member (7) can also extend beyond the perimeter ofcap and wrap (17) around it to come in contact with the under side ofcap (2). All or a portion of mounting device (6) with composite member(7) co-cured directly to cap (2) can be vacuum bagged and cured in anoven at about atmospheric pressure or in an autoclave at greater thanatmospheric pressure, preferably at about 80 to 200 psia.

Mounting device (6) can be installed into channel (9) defined in core(11) of skin (12) of an aircraft wing (10). As shown in FIGS. 3 and 4,an example of channel (9) can be defined in core (11) and up to insidemold line (16 a) of outer skin (16) of aircraft wing (10). The thicknessof composite member (7) of mounting device (6) can correspond to thethickness of core (11) of skin (12) along the length of aircraft wing(10). Mounting device (6) can be secured into channel (9) by using abonding agent, other adhesive, or both. Mounting device (6) can also besecured into channel (9) by co-curing mounting device (6) to inside moldline (16 a) of outer skin (16) of aircraft wing (10). All or a portionof the assembly can be can be vacuum bagged and cured in an oven atabout atmospheric pressure or in an autoclave at greater thanatmospheric pressure, preferably at about 80 to 200 psia.

As shown in FIG. 4, in one example, composite member (7) of upper andlower mounting devices, (6 a) and (6 b), can be inserted into opposingchannels, (9 a) and (9 b), in place of, for example, core (11) of skin(12) of aircraft wing (10). All or a portion of the assembly can be canbe vacuum bagged and cured in an oven at about atmospheric pressure orin an autoclave at greater than atmospheric pressure, preferably atabout 80 to 200 psia. Spar web panel (13) can be inserted into the upperand lower slots, (5 a) and (5 b), of mounting devices, (6 a) and (6 b).In an example of mounting device (6) with pinch angles (4), spaced apartarms (3) can pinch or engage spar web panel (13). A bonding agent, otheradhesive, or both can be used to further secure spar web panel (13) tomounting device (6). A tight fit between spaced apart arms (3) and sparweb panel (13) is preferred in part because it can allow the use of ahigher strength adhesive between spaced apart arms (3) and spar webpanel (13). If a bonding agent, other adhesives, or both (18) are usedto secure spar web panel (13) to mounting device (6), any excess canseep out holes (8), thereby allowing for slot (5) to be filled, as shownin FIG. 5. A paste adhesive is preferred to fill any gaps between thecap (2) and the spar web panel (13) in the slot (5). All or a portion ofthe assembly, of the upper and lower mounting devices, (6 a) and (6 b),in opposing channels, (9 a) and (9 b), with spar web panel installed, ofaircraft wing (10) can also be vacuum bagged and cured in an oven atabout atmospheric pressure or in an autoclave at greater thanatmospheric pressure, preferably at about 80 to 200 psia.

In the example shown in FIG. 4, mounting device (6) can comprisecomposite member (7) constructed of a graphite fiber epoxy laminatecomposite material and pi chord (1) constructed of a titanium alloy. Onebenefit of such an example is increased stiffness in aircraft wing (10).

In the example shown in FIGS. 3, 4, and 6, spar web panel (13) can betrimmed to the taper of wing (10) and inserted into both the upper andlower slots, (5 a) and (5 b), of mounting devices, (6 a) and (6 b), fromwing root (14) to tip (15) of the wing (10). Spar (13) can be secured tothe mounting devices, (6 a) and (6 b), as described above. All or aportion of the assembly can be vacuum bagged and cured in an oven atabout atmospheric pressure or in an autoclave at greater thanatmospheric pressure, preferably at about 80 to 200 psia.

One benefit of the example shown in FIGS. 3, 4, and 6 is that edges ofspar web (13) do not need to abut cap (2) along the length of mountingdevice (6) to be adequately secured within arms (3). Such an example,therefore, can allow for a dimensional mismatch between spar web panel(13) and the underside of caps (2) of two opposing mounting devices, (6a) and (6 b). Benefits of such an example include but are not limited tosimplified assembly and reduced incidence of “pull-in” or surface waveof outer mold line (16 b) of outer skin (16) of wing (10). The laterbeing a contributor to undesirable turbulent flow and drag over thesurface of the wing.

The invention is not limited to use as a mounting device in an aircraftwing. As shown in FIGS. 5, 7, and 8 the invention can serve as amounting device in other structural members of an aircraft. Otherstructural members contemplated by the invention include but are notlimited to a fuselage, fuselage skin, fin, or aileron of an aircraft ora floor or ceiling within an aircraft. In addition, the invention is notlimited to a mounting device constructed of a composite member (7) and ametal securement member that is configured in the form of a pi chord(1). As shown in FIG. 7, one example of the mounting device can beconstructed of a composite member (7) and a metal securement memberconfigured in the form of an I beam (19). As shown in FIG. 8, oneexample of the mounting device (6) can be constructed of a compositemember (7) and a metal securement member configured in the form of a Tchord (20). Other examples of the mounting device include but are notlimited to those constructed of a composite member and a metalsecurement member configured in the form of a J chord, sinewave beam,and F beam in either straight or curved sections (not shown). Inaddition, various methods known to those skilled in the art can be usedto mount or secure structures to the mounting device (6). These methodsinclude but are not limited to those described above and bolting,riveting, welding, curing, and clamping.

1. A method of installing a mounting device in a structural member of an aircraft, comprising the steps of: assembling a metal securement member by: holding unassembled metal sheets together in a constrained position so that the unassembled metal sheets do not move relative to one another; welding the unassembled metal sheets together with the metal sheets fixedly held in the constrained position so that the metal sheets do not move relative to one another during the welding; and annealing the welded together metal sheets while fixedly maintaining the welded metal sheets in the constrained position so that the welded metal sheets to not move relative to one another during the annealing; positioning a mounting device within a channel formed in the structural member, wherein the mounting device comprises the assembled metal securement member and a composite member attached to the metal securement member; and attaching the composite member to the structural member within the channel.
 2. The method of claim 1 wherein the assembled metal securement member comprises one of a pi chord, T chord, J chord, I beam, sinewave beam, or F beam.
 3. The method of claim 2 wherein the assembled metal securement member has a pi chord shape.
 4. The method of claim 2 wherein the assembled metal securement member has a T chord shape.
 5. The method of claim 2 wherein the assembled metal securement member has a J chord shape.
 6. The method of claim 2 wherein the assembled metal securement member has an I beam shape.
 7. The method of claim 2 wherein the assembled metal securement member has a sinewave beam shape.
 8. The method of claim 2 wherein the assembled metal securement member has a F beam shape.
 9. The method of claim 1 wherein the assembled metal securement member comprises: a cap member; and two spaced apart arms connected to and extending from a first surface of the cap member; wherein the composite member is attached to a second surface of the cap member, opposite the first surface.
 10. The method of claim 9 wherein the composite member is attached along an entire length and along an entire width of the second surface of the cap member.
 11. The method of claim 10 wherein the composite member, the cap member, and the channel all have identical lengths and identical widths.
 12. The method of claim 9 wherein the composite member extends beyond a perimeter of the cap member.
 13. The method of claim 9 further comprising bonding a spar to and between the two spaced apart arms with an adhesive so that excess of the adhesive seeps out holes, disposed along lengths of the two spaced apart arms, from between the two spaced apart arms to outside of the two spaced apart arms.
 14. The method of claim 1 wherein the structural member comprises at least one of a wing, wing skin, fuselage, fuselage skin, fin, aileron, floor, or ceiling.
 15. The method of claim 1 further comprising: positioning a second mounting device within a second channel formed in the structural member, wherein the second mounting device comprises a second assembled metal securement member and a second composite member attached to the assembled metal securement member; and attaching the second composite member to the structural member within the second channel.
 16. The method of claim 15 wherein the structural member is an aircraft wing.
 17. The method of claim 16 wherein the mounting device within the channel is positioned opposing the second mounting device in the second channel.
 18. The method of claim 17 further comprising attaching a spar to the mounting device and the second mounting device.
 19. The method of claim 18 wherein the spar attaches to the mounting device and the second mounting device over substantially an entire length of the wing.
 20. An aircraft wing manufactured according the method of claim
 1. 21. An aircraft manufactured according to the method of claim
 1. 22. The method of claim 1 wherein the metal securement member is made of a titanium alloy.
 23. The method of claim 1 wherein the composite member is made of a graphite fiber epoxy laminate.
 24. The method of claim 23 wherein the graphite fiber epoxy laminate comprises a lay-up of 70 to 80% zero degree fibers and 20 to 30% forty-five degree fibers.
 25. The method of claim 1 wherein the attaching comprises abutting and adhering the composite member to a bottom surface of the channel and to two opposed side surfaces of the channel.
 26. The method of claim 1 further comprising vacuum bagging and curing the assembled metal securement member and the composite member in an oven at about atmospheric pressure or in an autoclave at greater than atmospheric pressure.
 27. A method of installing a mounting device in a structural member of an aircraft, comprising the steps of: positioning a mounting device within a channel formed in the structural member, wherein the mounting device comprises a metal securement member, comprising a cap member and two spaced apart arms connected to and extending from a first surface of the cap member, and a composite member attached to a second surface of the cap member opposite the first surface; attaching the composite member to the structural member within the channel; and bonding a spar to and between the two spaced apart arms with an adhesive so that excess of the adhesive seeps out holes, disposed along lengths of the two spaced apart arms, from between the two spaced apart arms to outside of the two spaced apart arms. 